The present disclosure relates to a gas turbine engine combustor and, more particularly, to a liner panel therefor.
Gas turbine engines, such as those that power modern commercial and military aircraft, generally include a compressor section to pressurize an airflow, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases.
Among the engine components, relatively high temperatures are observed in the combustor section such that cooling airflow is provided to meet desired service life requirements. The combustor section typically includes an annular combustion chamber formed by an inner and outer wall assembly. Each wall assembly includes a support shell lined with heat shields, which are often referred to as liner panels. The liner panels may be segmented to accommodate thermal growth in operation and for other considerations. The combustor liner panels include a hot side exposed to the gas path. The opposite, or cold side, has features such as threaded studs to mount the liner panel to the support shell, and a perimeter rail that contacts the inner surface of the support shells.
The liner panel perimeter rail includes a forward rail that forms the forward, or upstream, edge of the panel, an aft rail that forms the aft, or downstream, edge of the liner panel, and longitudinal side rails that connect the forward and aft rails.
The liner panels extend over an arc in a conical or cylindrical array and axially interface in regions where the combustor geometry transitions (e.g., diverges or converges). This interface may contribute to durability and flow path concerns where forward and aft, as well as circumferentially adjacent, panels abut. These interfaces may be prone to steps between panels, dead regions, cooling challenges and adverse local aerodynamics.